Internal cooling cavity with trip strips

ABSTRACT

An airfoil is provided. The airfoil may comprise a cross over, an impingement chamber in fluid communication with the cross over, and a first trip strip disposed on a first surface of the impingement chamber. A cooling system is also provided. The cooling system may comprise an impingement chamber, a first trip strip on a first surface of the impingement chamber, and a second trip strip on a second surface of the impingement chamber. An internally cooled engine part is further provided. The internally cooled part may comprise a cross over and an impingement chamber in fluid communication with the cross over. The cross over may be configured to direct air towards a first surface of the impingement chamber. A first trip strip may be disposed on the first surface of the impingement chamber.

GOVERNMENT LICENSE RIGHTS

This disclosure was made with government support under contract No.N00019-12-D-0002 awarded by the United States Navy. The government hascertain rights in the disclosure.

FIELD OF INVENTION

The present disclosure relates to gas turbine engines, and, morespecifically, to flow guides for air and/or coolant flowing throughimpingement cavities.

BACKGROUND

Turbine airfoils or outer air seals operate in an environment where thegas temperatures often exceed the thermal capacity of materials in theengine. These parts may rely on cooling features to protect againstdamage. Cooling air from the compressor can be routed to provideinternal convection cooling within the airfoils. However, more coolingair bled from the compressor and used for cooling means less gas isavailable for work extraction. Thus, engine efficiency may be reduced ifhigher amounts of cooling air are consumed. As demands increase forhigher thrust and/or efficiency, the turbine inlet temperatures areincreased while the gas allocated for cooling is reduced.

Some components may implement air cooling systems with a series ofinternal cavities to cool a part. In some instances, the airrecirculates in an uncontrolled pattern before being bled off intoanother region of the part. The erratic air recirculation patterns maylimit the efficacy of internal flow cooling systems.

SUMMARY

An airfoil comprises a cross over, an impingement chamber in fluidcommunication with the cross over, and a first trip strip disposed on afirst surface of the impingement chamber.

In various embodiments, a second trip strip is disposed on a secondsurface of the impingement chamber. The first trip strip may have afirst geometry and the second trip strip may have a second geometrydifferent from the first geometry. The first trip strip and the secondtrip strip may be configured to direct air flow in a vortex motion. Anexit passage may be in fluid communication with the first trip strip andthe second trip strip. The first trip strip may be configured to directa first portion of air flow into the exit passage and a second portionof the air flow to the second trip strip. The first trip strip maycomprise a v-shaped geometry. The first trip strip may also comprise atleast one of a circular, elliptical, wave, or linear geometry.

A cooling system comprises an impingement chamber, a first trip strip ona first surface of the impingement chamber, and a second trip strip on asecond surface of the impingement chamber.

In various embodiments, a channel may be in fluid communication with theimpingement chamber, wherein the channel is configured to direct coolingfluid onto the first trip strip and the first surface of the impingementchamber. The first trip strip may comprise a v-shaped geometry. Thefirst trip strip may also comprise at least one of a circular,elliptical, wave, or linear geometry. An exit passage may be in fluidcommunication with the first trip strip and the second trip strip. Thefirst trip strip may be configured to direct a first portion of coolingfluid into the exit passage and a second portion of the cooling fluid tothe second trip strip. The first trip strip may have a first geometryand the second trip strip may have a second geometry different from thefirst geometry. The first trip strip and the second trip strip may beconfigured to direct cooling fluid in a vortex motion in the impingementchamber.

An internally cooled engine part may comprise a cross over and animpingement chamber in fluid communication with the cross over. Thecross over may be configured to direct air towards a first surface ofthe impingement chamber. A first trip strip may be disposed on the firstsurface of the impingement chamber.

In various embodiments, a second trip strip may be disposed on a secondsurface of the impingement chamber that is opposite the first surface.The first trip strip and the second trip strip may be configured todirect the air in a vortex motion in the impingement chamber. The firsttrip strip may comprise a v-shaped geometry.

The foregoing features and elements may be combined in variouscombinations without exclusivity, unless expressly indicated otherwise.These features and elements as well as the operation thereof will becomemore apparent in light of the following description and the accompanyingdrawings. It should be understood, however, the following descriptionand drawings are intended to be exemplary in nature and non-limiting.

BRIEF DESCRIPTION OF THE DRAWINGS

The subject matter of the present disclosure is particularly pointed outand distinctly claimed in the concluding portion of the specification. Amore complete understanding of the present disclosure, however, may bestbe obtained by referring to the detailed description and claims whenconsidered in connection with the figures, wherein like numerals denotelike elements.

FIG. 1 illustrates an exemplary embodiment of a gas-turbine engine, inaccordance with various embodiments;

FIG. 2 illustrates an airfoil with internal cavities for coolant flow,in accordance with various embodiments;

FIG. 3 illustrates a cross-sectional view of an airfoil with internalcavities for coolant flow, in accordance with various embodiments;

FIG. 4 illustrates a cross-sectional view of an internal cavity in anairfoil with guide features to direct cooling flow, in accordance withvarious embodiments;

FIG. 5 illustrates an internal cavity in an airfoil with guide featuresto direct cooling flow from impingement cavities, in accordance withvarious embodiments;

FIG. 6 illustrates the flow pattern over a suction-side surface of animpingement chamber, in accordance with various embodiments;

FIG. 7 illustrates a flow pattern over guide features on a pressure-sidesurface of an impingement chamber, in accordance with variousembodiments;

FIG. 8 illustrates various trip strip geometries on a pressure side ofan impingement chamber, in accordance with various embodiments; and

FIG. 9 illustrates various trip strip geometries on a suction side of animpingement chamber, in accordance with various embodiments.

DETAILED DESCRIPTION

The detailed description of exemplary embodiments herein makes referenceto the accompanying drawings, which show exemplary embodiments by way ofillustration. While these exemplary embodiments are described insufficient detail to enable those skilled in the art to practice theexemplary embodiments of the disclosure, it should be understood thatother embodiments may be realized and that logical changes andadaptations in design and construction may be made in accordance withthis disclosure and the teachings herein. Thus, the detailed descriptionherein is presented for purposes of illustration only and notlimitation. The scope of the disclosure is defined by the appendedclaims. For example, the steps recited in any of the method or processdescriptions may be executed in any order and are not necessarilylimited to the order presented.

Furthermore, any reference to singular includes plural embodiments, andany reference to more than one component or step may include a singularembodiment or step. Also, any reference to attached, fixed, connected orthe like may include permanent, removable, temporary, partial, fulland/or any other possible attachment option. Additionally, any referenceto without contact (or similar phrases) may also include reduced contactor minimal contact. Surface shading lines may be used throughout thefigures to denote different parts but not necessarily to denote the sameor different materials.

As used herein, “aft” refers to the direction associated with the tail(e.g., the back end) of an aircraft, or generally, to the direction ofexhaust of the gas turbine. As used herein, “forward” refers to thedirection associated with the nose (e.g., the front end) of an aircraft,or generally, to the direction of flight or motion.

As used herein, “distal” refers to the direction radially outward, orgenerally, away from the axis of rotation of a turbine engine. As usedherein, “proximal” refers to a direction radially inward, or generally,towards the axis of rotation of a turbine engine.

In various embodiments and with reference to FIG. 1, a gas turbineengine 20 is provided. Gas turbine engine 20 may be a two-spool turbofanthat generally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mayinclude, for example, an augmentor section among other systems orfeatures. In operation, fan section 22 can drive coolant (e.g., air)along a bypass flow-path B while compressor section 24 can drive coolantalong a core flow-path C for compression and communication intocombustor section 26 then expansion through turbine section 28. Althoughdepicted as a turbofan gas turbine engine 20 herein, it should beunderstood that the concepts described herein are not limited to usewith turbofans as the teachings may be applied to other types of turbineengines including three-spool architectures.

Gas turbine engine 20 may generally comprise a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A-A′ relative to an engine static structure 36 viaseveral bearing systems 38, 38-1, and 38-2. It should be understood thatvarious bearing systems 38 at various locations may alternatively oradditionally be provided, including for example, bearing system 38,bearing system 38-1, and bearing system 38-2.

Low speed spool 30 may generally comprise an inner shaft 40 thatinterconnects a fan 42, a low-pressure compressor 44 and a low-pressureturbine 46. Inner shaft 40 may be connected to fan 42 through a gearedarchitecture 48 that can drive fan 42 at a lower speed than low speedspool 30. Geared architecture 48 may comprise a gear assembly 60enclosed within a gear housing 62. Gear assembly 60 couples inner shaft40 to a rotating fan structure. High speed spool 32 may comprise anouter shaft 50 that interconnects a high-pressure compressor 52 andhigh-pressure turbine 54. Airfoils 55 of high-pressure turbine mayrotate about the engine central longitudinal axis A-A′. Airfoil 55 maybe an internally cooled component of gas turbine engine 20. Trip stripsmay be located in internal cooling cavities of internally cooled engineparts, as detailed further below. Internally cooled engine parts may bediscussed in the present disclosure in terms of airfoils. However, thepresent disclosure applies to any internally cooled engine part (e.g.,blade outer air seals, airfoil platforms, combustor components, or anyother internally cooled component in a gas turbine engine).

A combustor 56 may be located between high-pressure compressor 52 andhigh-pressure turbine 54. Mid-turbine frame 57 may support one or morebearing systems 38 in turbine section 28. Inner shaft 40 and outer shaft50 may be concentric and rotate via bearing systems 38 about the enginecentral longitudinal axis A-A′, which is collinear with theirlongitudinal axes. As used herein, a “high-pressure” compressor orturbine experiences a higher pressure than a corresponding“low-pressure” compressor or turbine.

The core airflow C may be compressed by low-pressure compressor 44 thenhigh-pressure compressor 52, mixed and burned with fuel in combustor 56,then expanded over high-pressure turbine 54 and low-pressure turbine 46.Mid-turbine frame 57 includes airfoils 59, which are in the core airflowpath. Turbines 46, 54 rotationally drive the respective low speed spool30 and high speed spool 32 in response to the expansion.

Gas turbine engine 20 may be, for example, a high-bypass ratio gearedaircraft engine. In various embodiments, the bypass ratio of gas turbineengine 20 may be greater than about six (6). In various embodiments, thebypass ratio of gas turbine engine 20 may be greater than ten (10). Invarious embodiments, geared architecture 48 may be an epicyclic geartrain, such as a star gear system (sun gear in meshing engagement with aplurality of star gears supported by a carrier and in meshing engagementwith a ring gear) or other gear system. Geared architecture 48 may havea gear reduction ratio of greater than about 2.3 and low-pressureturbine 46 may have a pressure ratio that is greater than about five(5). In various embodiments, the bypass ratio of gas turbine engine 20is greater than about ten (10:1). In various embodiments, the diameterof fan 42 may be significantly larger than that of the low-pressurecompressor 44. Low-pressure turbine 46 pressure ratio may be measuredprior to inlet of low-pressure turbine 46 as related to the pressure atthe outlet of low-pressure turbine 46 prior to an exhaust nozzle. Itshould be understood, however, that the above parameters are exemplaryof various embodiments of a suitable geared architecture engine and thatthe present disclosure contemplates other turbine engines includingdirect drive turbofans.

With reference to FIGS. 2 and 3, an airfoil 200 with internal cavities214 for carrying coolant flow (e.g., air flow) is shown according tovarious embodiments. Although an airfoil is shown, the presentdisclosure applies to any internally cooled part (e.g., blade outer airseals, airfoil platforms, combustor components, etc.). Airfoil 200 maycomprise leading edge 202 and trailing edge 208. Air flowing through agas turbine engine may first contact leading edge 202. Air may flowalong suction side 204 and/or pressure side 206 and leave airfoil attrailing edge 208. Airfoil 200 may include a blade platform 210 and anattachment root 212. Airfoil 200 is depicted as cutaway to illustrateinternal cavities 214 defined by internal walls 216. Internal cavities214 may be located throughout airfoil 200 and may provide internalcooling for airfoil 200.

With reference to FIG. 4, a cross-sectional view of an internal cavity214 in an airfoil 200 with trip strips 236 and trip strips 238 to directcooling flow, in accordance with various embodiments. Internal cavities214 may include an aft impingement chamber 230 containing trip strips236 disposed on an internal suction-side surface 237. Aft impingementchamber 230 may also contain trip strips 238 disposed on an internalpressure-side surface. A cross over 232 may be a narrow channeldirecting flow 234 into aft impingement chamber 230 where flow 234impinges on internal suction-side surface 237. Cross over 232 may directflow 234 as a jet into internal suction-side surface 237. In thatregard, cross over 232 may be oriented at an angle relative to internalsuction-side surface 237.

In various embodiments, flow 234 may eject from cross over 232 andcontact internal suction-side surface 237 and trip strips 236. Tripstrips 236 may direct flow 235 within aft impingement chamber 230 alonginternal suction-side surface 237. In that regard, trip strips 236 mayhave a tendency to prevent flow from moving radially outward as airfoil200 rotates (as airfoil 55 from FIG. 1 rotates about engine centrallongitudinal axis A-A′). Trip strips 236 may also provide increasedsurface area along internal suction-side surface 237 to improve heattransfer between the surface of airfoil 200 and the coolant making upflow 235. Flow 242 may then be ejected from aft impingement chamber 230by way of exit passage 240 before being ejected from airfoil 200.

With reference to FIG. 5, aft impingement chamber 230 with trip strips236 and trip strips 238 to direct cooling flow is shown in athree-dimensional view, in accordance with various embodiments. Internalsuction-side surface 237 is shown opposite internal pressure-sidesurface 239. Trip strips 236 are shown on internal suction-side surface237. Trip strips 238 are disposed on internal pressure-side surface 239,which defines in the opposite side of aft impingement chamber 230 frominternal suction-side surface 237. Cross overs 232 are shown as circularpassages directed towards internal suction-side surface 237. Althoughtrip strips 236 and trip strips 238 are shown with a v-shaped geometry,other geometries may also improve air flow characteristics and heattransfer within aft impingement chamber 230, as discussed further below.Similarly, although trip strips 236 and trip strips 238 are shown in aftimpingement chamber 230 of airfoil 200, trip strips may be deployed inother cooling cavities in airfoil 200. In fact, any aircraft part usingor comprising a part of an internal impingement chamber to providecooling may include trip strips, for example, a blade outer air seal.

With reference to FIGS. 6 and 7, the cooling flow pattern within aftimpingement chamber 230 is shown as viewed through internal suction-sidesurface 237, in accordance with various embodiments. Cooling flow flowsout from cross overs 232 and contacts internal suction-side surface 237on forward end 250, which is adjacent to the cross overs. Cooling airflowing into aft impingement chamber 230 initially spreads and thenfollows roughly along the contour of trip strips 236 and across internalsuction-side surface 237 to aft end 252. Trip strips 236 may act as aturbulator as well as a guide structure to air flow in aft impingementchamber 230. Trip strips 236 may also provide increased surface area toincrease heat transfer between internal suction-side surface 237 and thecooling air. Part of the cooling air flow may then exit through exitpassage 240 while the remainder of the cooling air flow is directed intointernal pressure-side surface 239.

In various embodiments, cooling air flow may contact internalpressure-side surface 239 at aft end 254 and be directed forward alonginternal pressure-side surface 239 in a forward direction. The coolingair flow may be directed guided roughly along trip strips 238 towardsforward end 256 of internal pressure-side surface 239. Cooling flow atforward end 256 may be directed towards the forward end 250 of internalsuction-side surface 237. The cooling flow may then mix with the coolantentering impingement chamber (i.e., flow 234) through cross overs 232and begin the cycle again. As shown in FIG. 4, flow 235 may travel in aswirling whirlpool or vortex motion within aft impingement chamber 230.

With reference to FIGS. 8 and 9, various trip strip geometries areshown. Trip strips 260 and 280 include wave geometry depicted foropposing surfaces in an impingement chamber. Trip strips 262 and tripstrips 282 include circular-platform geometries depicted for opposingsurface in an impingement chamber. Trip strips 266 and trip strips 286include elliptical-platform geometries depicted for opposing surface inan impingement chamber. Trip strips 264, trip strips 268, and tripsstrips 270 include linear geometries in various orientations, as do tripstrips 284, trip strips 288, and trip strips 290. Each geometricvariation of the trips strips may be used in conjunction with othergeometries on the same surfaces. For example, a surface may have a mixof trip strips 282 with a wave geometry and trip strips 284 with alinear geometry. Similarly, trips strips on opposing surfaces may vary.For example, trip strips 260 having wave geometry may be on a firstsurface of an impingement chamber while trip strips 282 havingcircular-platform geometry may be on the opposite surface. Furthermore,this disclosure contemplates any trip strip geometry being deployed inan impingement chamber.

Benefits and other advantages have been described herein with regard tospecific embodiments. Furthermore, the connecting lines shown in thevarious figures contained herein are intended to represent exemplaryfunctional relationships and/or physical couplings between the variouselements. It should be noted that many alternative or additionalfunctional relationships or physical connections may be present in apractical system. However, the benefits, advantages, and any elementsthat may cause any benefit or advantage to occur or become morepronounced are not to be construed as critical, required, or essentialfeatures or elements of the disclosure. The scope of the disclosure isaccordingly to be limited by nothing other than the appended claims, inwhich reference to an element in the singular is not intended to mean“one and only one” unless explicitly so stated, but rather “one ormore.” Moreover, where a phrase similar to “at least one of A, B, or C”is used in the claims, it is intended that the phrase be interpreted tomean that A alone may be present in an embodiment, B alone may bepresent in an embodiment, C alone may be present in an embodiment, orthat any combination of the elements A, B and C may be present in asingle embodiment; for example, A and B, A and C, B and C, or A and Band C.

Systems, methods and apparatus are provided herein. In the detaileddescription herein, references to “various embodiments”, “oneembodiment”, “an embodiment”, “an example embodiment”, etc., indicatethat the embodiment described may include a particular feature,structure, or characteristic, but every embodiment may not necessarilyinclude the particular feature, structure, or characteristic. Moreover,such phrases are not necessarily referring to the same embodiment.Further, when a particular feature, structure, or characteristic isdescribed in connection with an embodiment, it is submitted that it iswithin the knowledge of one skilled in the art to affect such feature,structure, or characteristic in connection with other embodimentswhether or not explicitly described. After reading the description, itwill be apparent to one skilled in the relevant art(s) how to implementthe disclosure in alternative embodiments.

Furthermore, no element, component, or method step in the presentdisclosure is intended to be dedicated to the public regardless ofwhether the element, component, or method step is explicitly recited inthe claims. No claim element herein is to be construed under theprovisions of 35 U.S.C. 112(f), unless the element is expressly recitedusing the phrase “means for.” As used herein, the terms “comprises”,“comprising”, or any other variation thereof, are intended to cover anon-exclusive inclusion, such that a process, method, article, orapparatus that comprises a list of elements does not include only thoseelements but may include other elements not expressly listed or inherentto such process, method, article, or apparatus.

What is claimed is:
 1. An airfoil, comprising: a cross over; animpingement chamber in fluid communication with the cross over; and afirst trip strip disposed on a first surface of the impingement chamber.2. The airfoil of claim 1, further including a second trip stripdisposed on a second surface of the impingement chamber.
 3. The airfoilof claim 2, wherein the first trip strip has a first geometry and thesecond trip strip has a second geometry different from the firstgeometry.
 4. The airfoil of claim 2, wherein the first trip strip andthe second trip strip are configured to direct an air flow in a vortexmotion.
 5. The airfoil of claim 2, further comprising an exit passage influid communication with the first trip strip and the second trip strip.6. The airfoil of claim 5, wherein the first trip strip is configured todirect a first portion of air flow into the exit passage and a secondportion of the air flow to the second trip strip.
 7. The airfoil ofclaim 2, wherein the first trip strip comprises a v-shaped geometry. 8.The airfoil of claim 2, wherein the first trip strip comprises at leastone of a circular, elliptical, wave, or linear geometry.
 9. A coolingsystem, comprising: an impingement chamber; a first trip strip on afirst surface of the impingement chamber; and a second trip strip on asecond surface of the impingement chamber.
 10. The cooling system ofclaim 9, further comprising a channel in fluid communication with theimpingement chamber, wherein the channel is configured to direct coolingfluid onto the first trip strip and the first surface of the impingementchamber.
 11. The cooling system of claim 9, wherein the first trip stripcomprises a v-shaped geometry.
 12. The cooling system of claim 9,wherein the first trip strip comprises at least one of a circular,elliptical, wave, or linear geometry.
 13. The cooling system of claim 9,further comprising an exit passage in fluid communication with the firsttrip strip and the second trip strip.
 14. The cooling system of claim13, wherein the first trip strip is configured to direct a first portionof a cooling fluid into the exit passage and a second portion of thecooling fluid to the second trip strip.
 15. The cooling system of claim9, wherein the first trip strip has a first geometry and the second tripstrip has a second geometry different from the first geometry.
 16. Thecooling system of claim 9, wherein the first trip strip and the secondtrip strip are configured to direct cooling fluid in a vortex motion inthe impingement chamber.
 17. An internally cooled engine part,comprising: a cross over; an impingement chamber in fluid communicationwith the cross over, wherein the cross over is configured to direct airtowards a first surface of the impingement chamber; and a first tripstrip disposed on the first surface of the impingement chamber.
 18. Theinternally cooled engine part of claim 17, further comprising a secondtrip strip disposed on a second surface of the impingement chamber, thesecond surface opposite the first surface.
 19. The internally cooledengine part of claim 18, wherein the first trip strip and the secondtrip strip are configured to direct the air in a vortex motion in theimpingement chamber.
 20. The cooling system of claim 17, wherein thefirst trip strip comprises a v-shaped geometry.